Gas turbine engines are commonly produced to include turbine nozzles, which accelerate and turn combustive gas flow toward the blades of a turbine rotor downstream of the nozzle. The turbine nozzle may have a generally annular or ring-shaped body including an inner endwall, an outer endwall circumscribing the inner endwall, and a series of circumferentially-spaced vanes extending between the inner and outer endwalls. The inner endwall, the outer endwall, and the vanes define a number of combustive gas flow paths through the turbine nozzle, which conduct hot combustive gas flow during operation of the gas turbine engine. While portions of the nozzle are exposed to combustive gas flow during engine operation, other portions of the turbine nozzle body and its associated mounting features are bathed in relatively cool airflow bled from a cold section of the engine and directed along an outer cooling flow path. In certain cases, undesired leakage can occur across the turbine nozzle interface between the outer cooling flow path and the core gas flow path. Such leakage can negatively affect the efficiency of the gas turbine engine, especially when smaller in size, and may increase the volume of airflow required for cooling purposes.
Leakage across the turbine nozzle mounting interfaces can be reduced through the usage of annular compression seals, such as flexible, pressure-activated metal seals. Such seals may be compressed between the mounting features of the turbine nozzle (e.g., rails extending radially from the opposing ends of the nozzle) and neighboring static structures within the engine. Temperature limitations may require that such compression seals are radially offset from the core gas flow path by a certain distance to reduce the operational temperatures to which the seals are exposed. The turbine nozzle rails may thus be elongated in a radial direction to allow such a radial offset between the compression seals and the core gas flow path. Unfortunately, this also has the effect of increasing temperature differentials that develop across the radially-elongated rails during engine operation, which may result in excessively high hoop stresses within the rails thereby hastening Thermomechanical Fatigue (TMF) and reducing the service lifespan of the turbine nozzle. TMF within the turbine nozzle rails may be alleviated through the formation of stress relief slots at strategic locations in the nozzle rail. The inclusion of stress relief slots in the nozzle rail may, however, permit an undesirably large amount of leakage across the turbine nozzle mounting interfaces thereby defeating the purpose of the compression seals or at least diminishing the effectiveness thereof.
It is thus desirable to provide embodiments of a turbine nozzle having stress relief slots formed at one or more circumferential locations in the radially-elongated rails or similar mounting features, which reduce TMF within the turbine nozzle while also minimizing leakage across the turbine nozzle mounting interfaces. More generally, it would be desirable to produce embodiments of a gas turbine engine component, such as a turbine nozzle or a combustor liner, including stress relief slots providing the above-noted benefits. Finally, it would be desirable to provide embodiments of a gas turbine engine employing such a gas turbine engine component, as well as methods for fabricating such a gas turbine engine component. Other desirable features and characteristics of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying Drawings and the foregoing Background.